Air breathing propulsion systems operating at high Mach numbers must use mixed-compression inlets in which the terminal shock, i.e. the beginning of subsonic flow, is located downstream of the inlet throat. The operation, and hence the control, of such propulsion systems critically depends on the streamwise position of the shock. Currently, there exists no reliable sensor of sufficient simplicity to provide terminal shock information to the engine control system.
Under laboratory conditions, it is possible to use a streamwise array of fast-response pressure transducers which are flush mounted on the inner wall of the flow channel. The transducers are placed to span the range of expected terminal shock locations. This laboratory system delivers an instantaneous pressure distribution, from which the terminal shock position is determined by locating the abrupt pressure rise associated with the terminal shock.
This system is damage prone and requires frequent calibration. The signal interpretation is both complicated and highly dependent upon flight conditions. It is therefore not well suited to the aerospace environment. Other techniques, such as optical methods and hot-film-based methods are also available, but they are only suitable for laboratory use.